During the past decade there has been growing interest within the space industry towards the development of small satellites. Small satellites are typically categorized as picosats (1 kg or less), nanosats (1-10 kg), microsats (10-100 kg) or minisats (100-500 kg) and range in size from softballs to refrigerators. The interest in these satellites is driven by the current constraints of traditional satellites and launch systems. As a result, there has been a significant effort to push satellite technology to smaller sizes and mass, which would enable small satellites to accomplish missions to complement the larger satellites. Examples of such missions include imaging, remote sensing, surveillance, disaster management, and blue force tracking. These missions are achieved by payloads which demand pointing capabilities from the satellites. This requires an attitude control system (ACS) with small actuators that can fit into the volume and mass constraints of small satellites.
The two major components of the ACS are the actuator and the control algorithm. Various types of actuators include the reaction wheel, magnetic rods, torque coils, thrusters, momentum wheels, and control moment gyroscope (CMG). CMGs rotate the angular momentum along a flywheel axis about a gimbal axis to produce a gyroscopic control torque as shown in FIG. 4(a). The output torque (gyroscopic torque) is amplified over the input torque required to rotate the gimbal axis (due to the satellite angular velocity) resulting in the well known torque amplification factor which allows for higher slew rates. This property of torque amplification as well as the fact that CMGs require minimal shaft power, permits the CMG to have a much higher torque per unit mass and unit power ratio than RWs.
More specifically, the CMG is a mechanism that produces torque by a combination of two motions—spinning a flywheel about an axis referred to as the flywheel axis and the rotation of the spinning flywheel about an axis perpendicular to flywheel axis referred to as the gimbal axis. The two main components of a gyroscope are the flywheel and the gimbal. The flywheel is a spinning rotor with inertia sufficient to provide the desired angular momentum; the gimbal is a pivot about which the flywheel assembly can be rotated. The magnitude of the gyroscopic torque produced is directly proportional to the inertia of the flywheel, the angular speed of the flywheel and the rate of rotation of the gimbal. In a CMG, the inertia of the flywheel and the speed of the flywheel are constant, and the torque output is controlled by changing the rotation rate of the gimbal. The direction of the torque produced is perpendicular to both the flywheel and the gimbal axes per the right hand rule. This torque acts on the satellite structure to change its attitude. A combination of gyroscopes is used to produce a net torque in the desired direction and magnitude. There are various combinations of gyroscopes that can be used depending upon the mission requirements (box configuration, inline configuration, roof top configuration, pyramidal configuration).
Apart from the gyroscopic torque produced by the CMG, there are other torques that arise from the motion of the flywheel and gimbal that contribute to the dynamics of the satellite:                Reaction torque due to friction in the flywheel bearings.        Reaction torque due to the acceleration of the gimbal; this torque depends on the angular acceleration and the inertia of the gimbal.        Reaction torque due to the friction of the gimbal bearings and slip ring.        
The motion to the flywheel and gimbal is provided by flywheel and gimbal motors. There are feedback devices (e.g., encoders and Hall-effect sensors) for sensing the angular speed and position. A slip ring is provided for continuous power supply to the flywheel motor for endless rotation of the gimbal. All these hardware are assembled together with structural components.
The output torque of the actuator is used to evaluate its performance. Certain kinds of actuators that use momentum from spinning wheels to generate torque are prone to disturbances due to misalignments and non-homogeneity of the wheels. This disturbance is termed as jitter. These actuators which contain wheels are evaluated for torque output and jitter as their performance metrics. Although there exist test beds that can evaluate the performance of large actuators, there has not been to date, an instrument to determine the performance of miniature actuators (<5 Nmm torque capacity).
ACSs are one of the most challenging spacecraft sub-systems for hardware performance verification and validation. Testing of ACSs pose many challenges as it requires simulation and control of spacecraft response to actuator inputs. Traditionally, this has been achieved using a motion platform that has three rotational degrees-of-freedom (dof) to emulate the spacecraft's attitude motion. ACS test beds developed to date for various institutions (e.g., Georgia Tech, NPS, Honeywell, AFRL to name a few) are typically for large spacecraft (>500 kg) and all, without exception, are suspension-based systems using air bearings1. Although these test beds have provided a means to validate 3-axis attitude control systems, they are limited in their testing capabilities and have several disadvantages. For example, ACS test beds using air bearings have limited range of motion about the pitch and roll axes (about ±30°)2, rendering the tests of continuous large angle maneuvers impossible. These test beds also have limits on their rotational rates as the centrifugal forces due to rotation have to be less than the dynamic capacity of the air bearing, i.e., these test beds limit the angular velocity of the simulated spacecraft thus preventing rapid retargeting maneuvers. Additionally, in order to nullify the bias torque due to gravity, these test beds require an additional dynamic mass balancing system which requires a separate control system3. The reactions due to the movement of the balancing masses induce unwanted disturbances to the spacecraft dynamics. There are additional effects/disturbances due to gravity sag associated with the size of these structures, viscous drag of the atmosphere, and the air draft from clean room blowers. Another major shortcoming of these test beds is that they cannot be used to test in an environmental chamber (thermal/vacuum) due to their complex nature and their dependency on air bearing. It is hence impossible to characterize the operational performance of the system under test in a representative space environment. FIG. 1(a) shows the schematic of a conventional test bed in which all the necessary hardware has to be integrated onto the air bearing platform. It should be noted that due to the platform's motion, an attitude determination system (ADS) is required onboard the air bearing test bed. ACS test beds for small satellites pose even harder challenges since they are more prone to viscous drags and other environmental disturbances due to their small inertia. Furthermore, it is difficult to incorporate a mass balancing system within the mass and inertia limits of these test beds. While there have been efforts in development of control strategies and actuators, there has been no transformative effort in the testing methods of attitude control system for the past fifty years1. There are currently no such test beds available for small satellites, especially for the pico (1 kg) and nano (10 kg) class.
Therefore, there exists a need for a system and method for assessing the performance of ACSs for satellites (e.g., pico and nano-satellites). It would be further advantageous for a system and method for assessing the performance of a CMG.